EP4331982B1 - Procédé de stabilisation d'une pale de rotor articulée - Google Patents

Procédé de stabilisation d'une pale de rotor articulée

Info

Publication number
EP4331982B1
EP4331982B1 EP23192923.3A EP23192923A EP4331982B1 EP 4331982 B1 EP4331982 B1 EP 4331982B1 EP 23192923 A EP23192923 A EP 23192923A EP 4331982 B1 EP4331982 B1 EP 4331982B1
Authority
EP
European Patent Office
Prior art keywords
rotor blade
thickness
spar
plies
aft
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
EP23192923.3A
Other languages
German (de)
English (en)
Other versions
EP4331982A1 (fr
Inventor
Joshua A. BREON
Eric S. Parsons
Christopher M. Sutton
William Paul Adams
Timothy James Conti
Jared Anapolle
Evan Medrano
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Lockheed Martin Corp
Original Assignee
Lockheed Corp
Lockheed Martin Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Lockheed Corp, Lockheed Martin Corp filed Critical Lockheed Corp
Publication of EP4331982A1 publication Critical patent/EP4331982A1/fr
Application granted granted Critical
Publication of EP4331982B1 publication Critical patent/EP4331982B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/32—Rotors
    • B64C27/46—Blades
    • B64C27/467—Aerodynamic features
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/32—Rotors
    • B64C27/46—Blades
    • B64C27/473—Constructional features

Definitions

  • the present invention relates generally to the field of rotor blade design for rotary wing aircraft.
  • Rotor blades for rotary wing aircraft such as helicopters, rotate to generate lift for vertical takeoff and flight. Reduction of the weight of rotor blades can reduce manufacturing cost and increase the payload of the aircraft.
  • US 5 332 362 A discloses a blade for aircraft rotary wings, with swept-back tip, wherein a variation in the length of the chord and the variation of the offset of the aerodynamic center are such that the leading edge and the trailing edge of the blade, along its longitudinal extent, exhibit no break, the overall aerodynamic center of said blade being substantially situated on the pitch variation axis.
  • GB 2 001 023 A discloses a torsionally compliant helicopter blade having its aerodynamic center axis positioned selectively forward of its center of gravity and elastic axes throughout a selected portion of the blade span so that, during operation, the blade will experience airloading which will induce two per rev torsional deflection of the blade to thereby improve rotor performance by reducing the total power required to drive the rotor.
  • a rotor blade for a rotary wing aircraft is provided.
  • the rotor blade includes a leading edge and a trailing edge defining a chordwise direction, the leading edge defining a forward end of the rotor blade and the trailing edge defining an aft end of the rotor blade.
  • the rotor blade has a center of gravity and an aerodynamic center along the chordwise direction, wherein the center of gravity is aft of the aerodynamic center.
  • the rotor blade further comprises a shear center that is forward of the aerodynamic center.
  • the rotor blade is dynamically stable.
  • the rotor blade further comprises a spar with a substantially oval cross section, the spar comprising an upper portion comprising an upper forward portion having a first thickness and an upper aft portion having a second thickness, and a lower portion comprising a lower forward portion having a third thickness and a lower aft portion having a fourth thickness, wherein the first thickness is larger than the second thickness, and the third thickness is larger than the fourth thickness.
  • the first thickness is substantially equal to the third thickness and the second thickness is substantially equal to the fourth thickness.
  • the rotor blade comprises a conduit between the spar and the leading edge of the rotor blade, the conduit being shaped to receive an electrical harness.
  • the upper forward portion comprises a first plurality of plies and the lower forward portion comprises a second plurality of plies.
  • each of the plies in the first and second pluralities of plies is above about 0.25 mm (0.010 inches) thick and below about 0.36 mm (0.014 inches thick.
  • the upper portion of the spar comprises an upper tapered portion positioned between the upper forward portion and the upper aft portion, and each of the plies in the first plurality of plies extends from an upper forward end of the upper forward portion to the upper tapered portion, and the lower portion of the spar comprises a lower tapered portion positioned between the lower forward portion and the lower aft portion, and each of the plies in the second plurality of plies extends from a lower forward end of the lower forward portion to the lower tapered portion.
  • the upper portion of the spar tapers from the first thickness to the second thickness within the upper tapered portion, and wherein the lower portion of the spar tapers from the third thickness to the fourth thickness within the lower tapered portion.
  • the upper tapered portion and the lower tapered portion each begin at a respective position between about 55% and about 65% of the distance to an aft end of the spar measured from a forward end of the spar, and the upper tapered portion and the lower tapered portion each end at a respective position between about 65% and about 75% of the distance to the aft end of the spar measured from the forward end of the spar.
  • rotor blades are designed with a center of gravity that is closer to the leading edge than the aerodynamic center (i.e., forward of the aerodynamic center) in order to provide dynamic stability to the blades and rotor system during flight.
  • Dynamic stability refers to the tendency of the blade to return to an equilibrium position once moved from that position by, for example, aerodynamic forces.
  • a dynamically unstable blade tends to oscillate with increasing amplitude once moved from an equilibrium position. This can result in undesirable vibration and/or harmonic motion of the blades that may affect or impair safe operation of the aircraft.
  • the center of gravity (e.g., the center of mass) refers to the average position of the mass of the rotor blade.
  • the aerodynamic center refers to the point at which the pitching moment of the blade does not substantially vary with the angle of attack of the blade.
  • typical blades are designed with a heavy weight at the leading edge, typically made of lead and/or tungsten.
  • a reduction in rotor blade weight can result in reduced manufacturing costs and higher aircraft payload. Accordingly, it would be advantageous to provide a lighter, dynamically stable rotor blade that does not require a heavy weight at the leading edge.
  • the exemplary non-limiting embodiments described herein provide a dynamically stable rotor blade with a center of gravity that is closer to the trailing edge than the aerodynamic center (i.e., aft of the aerodynamic center). This is achieved by designing the rotor blade with a shear center that is forward of the aerodynamic center. Shear center refers to the point at which force applied to the blade does not cause torsion.
  • the spar of the blade is thicker on the forward end than on the aft end such that the shear center is forward of the aerodynamic center.
  • the spar may be a multiple ply construction, with more plies on the forward end than on the aft end.
  • FIG. 1 an example of a vertical takeoff and landing (VTOL) rotary-wing aircraft 10 having a main rotor system 12, such as a helicopter, is illustrated.
  • the aircraft 10 includes an airframe 14 having an extending tail 16, which mounts a tail rotor system 18, such as an anti-torque system.
  • the main rotor system 12 is driven about an axis of rotation A via a main gearbox, illustrated schematically at T, by one or more engines, illustrated at E1-E3.
  • the main rotor system 12 includes a plurality of rotor blades 20 mounted to a rotor hub assembly H.
  • the rotor blades 20 may be rotor blades in accordance with the embodiments described herein.
  • helicopter configuration is illustrated and described herein, other configurations and/or aircraft may be utilized in connection with the concepts described herein.
  • the techniques according to the present disclosure may be implemented in a variety of aircraft, including a high speed compound rotary wing aircraft with supplemental translational thrust systems, a dual contra-rotating rotary aircraft, coaxial rotor system aircraft, a tilt-rotor aircraft, and a tilt-wing aircraft.
  • the rotor blade 100 includes a leading edge portion 102 including a leading edge 104 (e.g., a forward edge) that defines a forward end of the rotor blade 100, and a trailing edge portion 106 including a trailing edge 108 (e.g., an aft edge) that defines an aft end of the rotor blade 100.
  • the primary structural member is a spar 110, the cross section of which is roughly oval (e.g., substantially ovate) in shape.
  • the spar 110 also includes a leading edge 112 (e.g., a forward edge), and a trailing edge 114 (e.g., an aft edge).
  • the spar 110 extends from the root to the tip of the blade 100.
  • the spar 110 defines an internal cavity 111 that may be filled with a lightweight material such as foam, or may not be filled at all.
  • the outer surface of the rotor blade 100 includes an upper skin 116 and a lower skin 118 that are bonded to the spar 110 and extend to the trailing edge portion 106 to form the trailing edge 108, as well as a leading edge sheath 105 at the leading edge portion 102 that forms the leading edge 104.
  • the trailing edge portion 106 includes a lightweight core 122 positioned between the upper and lower skins 116, 118 and the trailing edge 214 of the spar, that may comprise, for example, a lightweight foam or lightweight honeycomb structure.
  • the leading edge portion 102 of the blade 100 includes a forward weight 120 positioned between the upper and lower skins 116, 118 and the leading edge of the spar 112. Because the internal cavity 111 of the spar 110 and the core 122 of the trailing edge portion 106 include lightweight material or no material at all, the forward weight 120 ensures that, chordwise, the center of gravity 124 is forward of the aerodynamic center 126 of the blade 100. This provides dynamic stability to the blade 100 and the rotor system during flight.
  • the forward weight 120 may be made from, for example, lead or tungsten, and may comprise as much as approximately half of the total weight of the rotor blade 100. For example, a conventional blade may have a total weight of approximately 250 lbs. and a forward weight of approximately 125 lbs.
  • the rotor blade 200 includes a leading edge portion 202 including a leading edge 204 (e.g., a forward edge) that defines a forward end of the rotor blade, and a trailing edge portion 206 including a trailing edge 208 (e.g., an aft edge) that defines an aft edge of the rotor blade.
  • the rotor blade 200 includes a spar 210, the cross section of which is roughly oval in shape.
  • the spar 210 also includes a leading edge 212 (e.g., a forward edge) defining a forward end of the spar 210, and a trailing edge 214 (e.g., an aft edge) defining an aft end of the spar.
  • the spar 210 begins at the inboard root of the rotor blade 200 and extends most of the way to the outboard tip of the rotor blade 200.
  • the spar 210 defines an internal cavity 211 that may be filled with a lightweight material such as foam, or may remain an empty space.
  • the outer surface of the rotor blade 200 includes an upper skin 216 and a lower skin 218 that are bonded to the spar 210 and extend to the trailing edge portion 206 to form the trailing edge 208, as well as a leading edge sheath 105 at the leading edge portion 202 that forms the leading edge 204.
  • the spar 210 and upper and lower skins 216, 218 may be made from titanium, aluminum, stainless, steel, carbon fiber, fiberglass, or any other appropriate material.
  • the spar 210 and/or the upper and lower skins 216, 218 may be made from a multifilament graphite having a typical filament tensile modulus of approximately 38 x 10 6 to approximately 42 x 10 6 psi (approximately 262 to approximately 290 GPa) and a typical filament tensile strength of approximately 660,000 psi (approximately 4452 MPa).
  • the trailing edge portion 206 includes a lightweight core 222 positioned between the upper and lower skins 216, 218 and the trailing edge 214 of the spar, that may be made from, for example, a lightweight foam or a lightweight honeycomb type structure.
  • the rotor blade 200 does not include a forward weight positioned between the spar 210 and the leading edge 204 of the rotor blade 200. Instead, the leading edge portion 202 of the blade 200 includes a cavity 220.
  • the cavity 220 may be filled with a lightweight material such as foam, may remain an empty space, or may include conduit for running electrical wiring to the outboard end of the rotor blade 200.
  • wire harness assemblies that can be implemented within a conduit in the cavity 220, as well as additional structural details of any electrical connections, can be found in U.S. Patent Application No. 17/485,192 filed September 24, 2021 .
  • the cavity 220 is configured to house a wire harness extending through an inner portion of a conduit, which is also housed in the cavity 220.
  • the cavity 220 extends from at least an inboard portion to at least an outboard portion of the rotor blade 200, and the conduit extends from the inboard to the outboard portion.
  • the wire harness is configured to electrically couple an electrical component at the outboard portion of the rotor blade to an electrical component at the inboard end of the rotor blade 200.
  • a filler block holds the conduit within the cavity 220.
  • the center of gravity 224 is aft of the aerodynamic center 226 such that, chordwise, the aerodynamic center 226 is between the center of gravity 224 and the sheet center 228. In a traditional rotor blade design, this would result in a dynamically unstable blade and rotor system. However, due to the design of the blade 200, dynamic stability is achieved because the shear center 228 of the blade 200 is forward of the aerodynamic center 226. The locations of the aerodynamic center 226, shear center 228, and center of gravity 224 are dependent on the design of the blade 220, including the design of the spar 210. In the embodiment shown in FIG.
  • the shear center 228 of the blade 200 is moved forward due primarily to the design of the spar 210.
  • the location of the shear center 228 is a function of the relative bending stiffness of the forward portion of the spar 210 compared to the aft section of the spar. If the forward portion manufactured to be stiffer than the aft portion, the shear center is biased in the forward direction.
  • the spar 210 includes an upper portion 239 and a lower portion 243.
  • the upper portion 239 includes an upper forward portion 240 and an upper aft portion 242.
  • the lower portion includes a lower forward portion 244 and a lower aft portion 246.
  • the spar 210 is manufactured such that the forward portions 240, 244 are stiffer than the aft portions 242, 246.
  • the upper and lower forward portion 240, 244 are each thicker than the upper and lower aft portions 242, 246.
  • the thickness of the upper portion 239 tapers prior to the trailing edge 214 at an upper tapered portion 241 from a first thickness T1 at the upper forward portion 240 to a second thickness T2 at the upper aft portion 242.
  • the thickness of the lower portion 243 tapers prior to the trailing edge 214 at a lower tapered portion 245 from a third thickness T3 at the lower forward portion 244 to a fourth thickness T4 at the lower aft portion 246.
  • the upper and lower tapered portions 241, 245 may begin approximately 55% to approximately 65% of the distance to the trailing edge 214, and may end approximately 65% to approximately 75% of the distance to the trailing edge 214.
  • the locations of the tapered portions 241, 245 may be different depending on the relative thicknesses and material stiffness properties of the forward portions 240, 244 compared to the aft portions 242, 246.
  • thickness T1 may be substantially equal to thickness T3.
  • thickness T2 may be substantially equal to thickness T4.
  • the upper portion 139 and the lower portion 141 of the conventional spar 110 each have a substantially consistent thickness.
  • the location of the shear center is not determinative of the dynamic stability of the blade 100 because the center of gravity 124 is forward of the aerodynamic center 126. While this results in a dynamically stable blade, it requires that the blade include the forward weight 120 in order to move the center of gravity 124 forward. Because the design of the spar 210 in the embodiment shown in FIGS. 3 and 4 moves the shear center 228 forward of the aerodynamic center 226, no forward weight is required.
  • the upper forward portion 240 includes a first plurality of bonded unidirectional plies 250 (e.g., layers), and the lower forward portion 244 includes a second plurality of unidirectional plies 252.
  • "Unidirectional" refers to the orientation of the majority of fibers in the layers. The majority of the fibers are oriented in a zero degree direction parallel to the feathering axis of the blade 200.
  • Each of the plies in the first and second pluralities of plies may be between about 1/100" thick to about 1/70" thick, inclusive, in some embodiments, e.g., about 1/80" thick or 1/90" thick.
  • the first plurality of plies 250 extend from an upper forward end 238 adjacent the leading edge 212 of the spar 210 to specific locations within the upper tapered portion 241.
  • the second plurality of plies 252 extend from a lower forward end 247 adjacent the leading edge 212 of the spar 210 to specific locations within the lower tapered portion 245.
  • the first plurality of plies includes a first ply pack 261, a second ply pack 262, a third ply pack 263, a fourth ply pack 264, and a fifth ply pack 265.
  • Each ply pack 261, 262, 263, 264, 265 includes a plurality of plies.
  • the first ply pack 261 is bonded to the underside of the upper skin 216.
  • Each successive ply pack is bonded to the previous ply pack forming a layered material extending towards the internal cavity 211.
  • the fourth ply pack 264 tapers and ends at a point 266 as it approaches the upper aft portion 242.
  • the third and fifth ply packs 263, 265 merge together at point 266.
  • the second ply pack 262 tapers and ends at a point 267 as it approaches the upper aft portion 242.
  • the first ply pack 261 and the combined third and fifth ply packs 263, 265 merge together at point 267.
  • Each of the ply packs 261-265 may be about 1/20" thick to about 1/7" thick, inclusive, e.g., about 1/18" or about 1/8" thick.
  • the thickness may be determined based on the desired shear center.
  • the ply pack is a multifilament material, such as a multifilament graphite.
  • the buildup of plies near the forward end of the spar 210 moves the shear center 228 forward until the shear center 228 is forward of the aerodynamic center 226.
  • Various embodiments of the rotor blade 200 may include more or fewer plies than as shown in FIG. 5 .

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Wind Motors (AREA)

Claims (9)

  1. Pale de rotor (200) pour un aéronef à voilure tournante, la pale de rotor comprenant :
    un bord d'attaque (204) et un bord de fuite (208) définissant une direction dans le sens de la corde, le bord d'attaque (204) définissant une extrémité avant de la pale de rotor et le bord de fuite (208) définissant une extrémité arrière de la pale de rotor,
    la pale de rotor (200) ayant un centre de gravité (224) et un centre aérodynamique (226) le long de la direction dans le sens de la corde, caractérisée en ce que le centre de gravité (224) est à l'arrière du centre aérodynamique (226) et la pale de rotor comprend également un centre de cisaillement (228) qui est à l'avant du centre aérodynamique (226).
  2. Pale de rotor selon la revendication 1, dans laquelle la pale de rotor est dynamiquement stable.
  3. Pale de rotor selon la revendication 1 ou 2, comprenant également un longeron (210) avec une section transversale sensiblement ovale, le longeron comprenant :
    une partie supérieure (239) comprenant une partie avant supérieure (240) ayant une première épaisseur et une partie arrière supérieure (242) ayant une deuxième épaisseur ; et
    une partie inférieure (243) comprenant une partie avant inférieure (244) ayant une troisième épaisseur et une partie arrière inférieure (246) ayant une quatrième épaisseur ;
    dans laquelle la première épaisseur est supérieure à la deuxième épaisseur, et la troisième épaisseur est supérieure à la quatrième épaisseur.
  4. Pale de rotor selon la revendication 3, dans laquelle la première épaisseur est sensiblement égale à la troisième épaisseur et la deuxième épaisseur est sensiblement égale à la quatrième épaisseur.
  5. Pale de rotor selon l'une ou l'autre des revendications 3 ou 4, dans laquelle la pale de rotor (200) comprend un conduit entre le longeron (210) et le bord d'attaque (204) de la pale de rotor, le conduit étant façonné pour recevoir un faisceau électrique.
  6. Pale de rotor selon l'une quelconque des revendications 3 à 5, dans laquelle la partie avant supérieure (240) comprend une première pluralité de plis (250) et la partie avant inférieure (244) comprend une deuxième pluralité de plis (252).
  7. Pale de rotor selon la revendication 6, dans laquelle chacun des plis des première et deuxième pluralités de plis (250, 252) a une épaisseur supérieure à environ 0,25 mm et inférieure à environ 0,36 mm.
  8. Pale de rotor selon l'une ou l'autre des revendications 6 ou 7, dans laquelle :
    la partie supérieure (239) du longeron comprend une partie conique supérieure (241) positionnée entre la partie avant supérieure (240) et la partie arrière supérieure (242), et chacun des plis de la première pluralité de plis (250) se prolonge d'une extrémité avant supérieure (238) de la partie avant supérieure (240) à la partie conique supérieure (241) ; et
    la partie inférieure (243) du longeron comprend une partie conique inférieure (245) positionnée entre la partie avant inférieure (244) et la partie arrière inférieure (246), et chacun des plis de la deuxième pluralité de plis (252) se prolonge d'une extrémité avant inférieure (247) de la partie avant inférieure (244) à la partie conique inférieure (245).
  9. Pale de rotor selon la revendication 8, dans laquelle la partie supérieure (239) du longeron s'effile de la première épaisseur à la deuxième épaisseur au sein de la partie conique supérieure (241), et dans laquelle la partie inférieure (243) du longeron s'effile de la troisième épaisseur à la quatrième épaisseur dans la partie conique inférieure (245).
EP23192923.3A 2022-08-31 2023-08-23 Procédé de stabilisation d'une pale de rotor articulée Active EP4331982B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US17/900,780 US12110104B2 (en) 2022-08-31 2022-08-31 Method of stabilizing articulated rotor blade

Publications (2)

Publication Number Publication Date
EP4331982A1 EP4331982A1 (fr) 2024-03-06
EP4331982B1 true EP4331982B1 (fr) 2026-01-21

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP23192923.3A Active EP4331982B1 (fr) 2022-08-31 2023-08-23 Procédé de stabilisation d'une pale de rotor articulée

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US (2) US12110104B2 (fr)
EP (1) EP4331982B1 (fr)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US12110104B2 (en) * 2022-08-31 2024-10-08 Lockheed Martin Corporation Method of stabilizing articulated rotor blade

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CA889178A (en) * 1971-12-28 J. Dutton Walter Helicopter blade with non-homogeneous structural spar
US3065799A (en) 1958-04-21 1962-11-27 Brackley Shaw Rotary wing aircraft
US3323597A (en) 1965-02-01 1967-06-06 United Aircraft Corp Bonded counterweight for blade of rotary wing aircraft
US3528753A (en) 1968-06-14 1970-09-15 United Aircraft Corp Helicopter blade with non-homogeneous structural spar
CA988906A (en) * 1973-06-04 1976-05-11 United Aircraft Corporation Helicopter blade
US4130377A (en) * 1977-07-13 1978-12-19 United Technologies Corporation Helicopter blade and rotor
FR2689852B1 (fr) * 1992-04-09 1994-06-17 Eurocopter France Pale pour voilure tournante d'aeronef, a extremite en fleche.
CA2198717C (fr) * 1994-08-31 2005-03-15 William C. Reinfelder Longeron composite renforce par fibres pour aeronef a voilure tournante et son procede de fabrication
US6986642B2 (en) * 2002-08-30 2006-01-17 Cartercopters, L.L.C. Extreme mu rotor
US9139297B1 (en) * 2010-09-09 2015-09-22 Groen Brothers Aviation, Inc. Rotor blade subsystems attachment
US10239604B2 (en) * 2016-05-21 2019-03-26 Bell Helicopter Textron Inc. Structurally biased proprotor blade assembly
US12110104B2 (en) * 2022-08-31 2024-10-08 Lockheed Martin Corporation Method of stabilizing articulated rotor blade

Also Published As

Publication number Publication date
US20250010983A1 (en) 2025-01-09
US12110104B2 (en) 2024-10-08
US20240067333A1 (en) 2024-02-29
EP4331982A1 (fr) 2024-03-06

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